General Electric J47-GE-17B and J47-GE-33 Engine Installation
Part 4: Engine Systems
Compiled by Kimble D. McCutcheon

 

Anti-Icing

The J47-GE-17B and -33 were completely anti-iced all-weather power plants. Those engine parts with a frontal area exposed to the inlet air stream required icing protection. Included were the island fairings, front frame struts, inlet guide vanes, and air screens. In addition, the aircraft manufacturer's bullet-nose also required anti-icing and provisions were made for the engine to supply hot air to the bullet-nose. The air screens had no ice protection. Their only protection arose from the fact that they could be retracted in flight. The retracting mechanism consisted of eight jackscrews connected in series by a flexible shaft that was driven by a 28 VDC motor. Maximum power requirement was 5 amps. The flexible shaft completely encircled the engine and each jack incorporated a friction clutch so that if any one section failed the others could still be actuated. The bullet-nose was anti-iced by compressor discharge air taken off a pad on the compressor rear frame, brought forward through Island No. 4, and then to the bullet-nose.

The island fairings were anti-iced by means of compressor discharge air that was bled off the bullet-nose line, into a manifold, and then to the island fairings. The air was fed into a cavity at the fairing leading edges and then discharged into the engine air stream through vents on the fairing sides. The air for the fairings and the bullet-nose could be turned off or on by means of a solenoid-operated pilot valve that operated the anti-icing air valve. The pilot valve switch was controlled from the cockpit. Some installations had a thermo-switch in the bullet-nose that was in series with the cockpit switch; it cut off the anti-icing system above 42°F. This thermo-switch was furnished by the airframe manufacturer.

The inlet guide vanes and front frame struts were anti-iced continuously by means of 12th stage air taken off at two points on the compressor casing 180° apart. For the guide vanes, the air was brought in through two elbows 180° apart into a manifold that discharged into the balance piston area through thirty-seven hollow inlet guide vanes.

Exhaust System

The engine exhaust system was composed of an exhaust reheat burner and a variable area nozzle that furnished takeoff, climb, and maximum speed condition thrust augmentation. This was accomplished by introducing additional fuel to burn with the oxygen present in the exhaust gas stream. The jet nozzle area was controlled to maintain turbine discharge temperature at its maximum safe limit for both dry and reheat operation.

The reheat process consisted of the following phases:
Phase 1 - Diffusion of turbine gases to lower velocities was accomplished in a section of increasing area following the turbine. The outer cone converged slowly and the inner cone converged rapidly to form this diffusing section.
Phase 2 - Fuel injection was carried out by two circumferential nozzle rows of 19 nozzles each that extended radially outward from the inner cone midway along its outer surface. For low reheat fuel flow only the downstream nozzles were used, but as the flow was increased both nozzle sets inject fuel. The nozzles themselves consisteded of short lengths of tubing in which holes were drilled to properly inject and distribute the fuel. Each circumferential nozzle row was manifolded inside the inner cone and fuel from the reheat flow divider was piped to the tailcone and then radially to the reheat manifold. The fuel lines did not pass through a support strut fairings, but were exposed to the exhaust gas stream.
Phase 3 - Fuel vaporization and mixing occurred immediately downstream of the fuel nozzles. The nozzle hole locations and sizes were selected to properly distribute the fuel in the hot gas stream. The high velocity gases readily separated each fuel stream into fine particles that were mixed and vaporized to form a combustible mixture.
Phase 4 - Ignition was initiated by a hot gas stream from the hot streak ignition jet. A double V-shaped ring flame holder was used to decrease the mixture burning velocity, thus allowing time for complete combustion and provided an efficient, self-sustaining burning process.
Phase 5 - Flame propagation and burning occurred aft of the pilot dome and flame holder. Flame propagating in the shape of a cone that fills the tailpipe at full augmentation. The reheat burner walls were insulated from the high reheat temperatures by a thin layer of turbine discharge gas and a ceramic coated liner in the reheat burner.

The variable area nozzle was mounted on the tailpipe end and was automatically controlled to provide a range of areas from 96% to 160% of the normal fixed area for this type of engine. The nozzle consisted of a fixed portion with two clamshell lids at the aft end, which were pivoted about a vertical axis perpendicular to the engine center line. These two clamshells were linked to the nozzle actuator that was mounted on the exterior at the bottom of the tailpipe. During reheat operation the nozzle was cooled by compressor discharge air.

Cooling System

The J47 incorporated several internal cooling features:
The turbine wheel was cooled by air bled from the 12th compressor stage to the turbine wheel aft face, and by air bled from the 8th stage to turbine wheel the forward side.
Ignitor plugs and cross-over tubes were cooled by air from the compressor discharge fed between the combustion liner and outer chamber. At the combustion liner aft end this secondary cooling air was metered through a slotted L seal, guided through the hollow nozzle diaphragm blades, and out into the exhaust stream along the shroud ring inner surface, thereby cooling the nozzle diaphragm blades and providing an air cooling blanket between the turbine wheel and shroud.
The variable nozzle and its actuator were cooled by compressor discharge cooling air supplied at 0.5 lbs/sec by two diametrically opposite ducts from the compressor rear frame through a cooling air control valve, to the nozzle aft end. The air operated control valve was controlled by a solenoid-operated pilot valve that allowed cooling air flow only during reheat. In the variable nozzle cooling air was manifolded into an annular chamber at the aft end and then escaped through drilled holes in the outer surface to the nozzle lids underside. The jet nozzle actuator was housed in a metal casing, and cooling air at atmospheric pressure was be supplied to the housing inlet.

 

Other Systems

The lubrication (lube) system provided an adequate clean oil supply at the correct temperature for lubricating bearings and gears in the accessory and PTO case, and for cooling and lubricating the engine bearings. The lube system was a recirculating positive displacement type composed of a three-element lube and scavenge pump, a double element aft frame scavenge pump, a single element auxiliary accessory gear case scavenge pump, an oil flow divider, an oil cooler and lube oil filter.

Oil from an airframe-mounted tank entered the main lube pump through Island No. 2. Pump element No. 1 sent oil out through Island No. 3 and aft through the oil filter to the compressor rear frame where the oil piping entered the engine and passed down to the lube oil flow divider. The flow divider functioned so that the primary system built up pressure rapidly at starting and the secondary cut in at 15 psi. This operational characteristic produced a high penetration jet at starting and a constant pressure lube system from idle to design rpm. In the flow divider split 13% of the oil to the primary lube piping and jets, and the other 87% to the secondary or main lube piping and jets. Oil from these jets lubricated bearings Nos. 2,3 and 4, and was then scavenged from the aft frame and sumps in the compressor rear frame by the two element scavenge pump. The discharge from this pump passes out through the compressor rear frame, entered the lube oil cooler and was then piped to a scavenge manifold on the No. 3 island. Although the lube pump No. 1 element supplied 2.5 gpm to the Nos. 2, 3, and 4 bearings, this oil was not equally distributed; the No. 2 bearing received 1.5 gpm, the No. 3 bearing 0.5 gpm and the No. 4 bearing 0.5 gpm.

The No. 2 element sent oil into passages in the main gear case, passing through the static leakage valve, a strainer below the pressure relief valve, and then to the bearings and gears in the gear case forward section. Branching from the same oil line, a section of piping supplied oil to the No. 1 bearing and the gears and bearings in the case rear. Oil was bled from this system and entered oil passages in the auxiliary gear case where two jets lubricated the gears from a central manifold. The auxiliary gear case was scavenged by a small scavenge pump and the main gear case by the scavenge element in the main lube pump. Both scavenge discharges were joined together and went through the No. 3 island, picking up the oil in the scavenge line from the aft frame and returning it to the lube tank. The No. 2 element pumped 2.5 gpm of which 1.5 gpm went to the accessory gear case and 1.0 gpm to the auxiliary gear case.

Total oil retention at various engine speeds was as follows:
Shutdown = 5.5 ± 0.5 lb
Idle rpm = 8.0 ± 1.0 lb
Military rpm = 13.0 ± 1.0 lb.

Fuel System. Tank mounted boost pumps supplied fuel under pressure through a common line to the main fuel system, through the main system shut-off valve and a line-mounted filter to the engine fuel pump, which was a two-element positive displacement type with each element on a separate shearable shaft. A differential pressure switch connected to each pump element's discharge side indicated a partial failure of either pump via a cockpit warning light. Each pump's discharge passed through an independent check valve into a common line supplying the control valve. The main fuel control valve was a flow-control type that regulated fuel flow to the engine as a function of an input shaft position by passing the excess fuel back to the fuel pump inlet. Fuel going to the engine passed through a manually actuated stopcock connected to the pilot's quadrant. The engine fuel then passes through an oil cooler to a flow divider and to the nozzles in each combustion chamber. The nozzles were shrouded to direct air across their tip, which reduced both carbon formation on the nozzles and engine smoking.

A mechanical bypassing type engine-driven overspeed governor was connected to the fuel line downstream of the fuel control valve. When an overspeed condition existed, the governor opened delivering back enough fuel back to the pump inlet to hold the engine to a safe overspeed.

An emergency fuel regulator was located in the bypass line from the main fuel control valve. If the main control system failed the emergency regulator took over fuel flow control with the aid of the stopcock and a valve bypass in the control valve unit.

Reheat System. A tank-mounted reheat fuel pump, which was driven by compressor discharge air and controlled by an electrically operated shutoff valve mounted in the air line, pumped fuel to a check valve and fuel filter furnished by the airframe manufacturer. The fuel then passed to the reheat fuel control valve, a flow-control type valve that regulated fuel flow to the engine as a function of an input shaft position, by passing excess fuel back into the main fuel tank. From the control valve the reheat fuel going to the engine was divided between two fuel manifolds by a flow divider. A check valve in the primary line, or the line that bypassed the flow divider, prevented tailpipe pressure from causing back flow in the fuel line.

The electrical system provided a way of starting the engine and provided an electrical power source. For starting, electrical power was usually supplied by an external source, normally a motor-generator set. In this case, the starter-generator functioned as a direct-current motor to accelerate the engine to a speed where the air flow and fuel pressure were such that combustion would occur upon ignition. Ignition was supplied by engine mounted vibrators and coils through opposite polarity spark plugs. After ignition occurred, sufficient energy became available from the hot gas to enable the turbine, in conjunction with the starter, to accelerate the engine to a minimum stable operating speed of approximately 2,000 rpm.

Starting was accomplished by first turning on the main engine power switch. This switch connected the starter control to the main 24 VDC bus. A momentary contact start switch energized the undercurrent relay that in turn held the starter contactor closed until the engine reached approximately 2,000 rpm. This enabled the starter-generator to assist the turbine during the initial engine acceleration.


 

On to Part 5

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